The present disclosure relates to a gas turbine engine and more particularly to a turbine vane cooling arrangement.
Gas turbine engines, such as those which power modern military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust. Downstream of the turbine section, an augmentor section, or “afterburner”, is operable to selectively increase the thrust. The increase in thrust is produced when fuel is injected into the core exhaust gases downstream of the turbine section and burned with the oxygen contained therein to generate a second combustion.
The turbine section typically includes alternating rows of turbine vanes and turbine blades. The turbine vanes are stationary and function to direct the hot combustion gases that exit the combustor section. Due to the relatively high temperatures of the combustion gases, various cooling techniques are employed to cool the turbine vanes and blades.
The vanes typically include a hollow airfoil section with a leading edge wall followed by a pressure side wall and a suction side wall that converge to form a trailing edge. The hollow airfoil section is typically cooled with bleed air from the compressor section. Among the various cooling techniques are convection, impingement, film cooling as well as radiation within and through the airfoil wall surfaces.
Further, cooling airflows are often passed thru the turbine vanes to cool radially inboard or outboard components and structures. Although effective, the multiple cooling schemes result in a relatively complex inner vane structure which may transfer heat from the airfoil wall surfaces to the pass thru air and reduce the cooling effectiveness thereof.